1. Field of the Invention
The present invention relates generally to gas turbine engine turbine blade tip cooling and, more specifically, to a turbine blade tip coated with thermal barrier.
2. Description of Related Art
A gas turbine engine turbine blades extract energy from hot combustion gas for powering the compressor and providing output power. Since the turbine blades are directly exposed to the hot combustion gas, they are typically provided with internal cooling circuits which channel a coolant, such as compressor bleed air, through the airfoil of the blade and through various film cooling holes around the surface thereof.
One type of airfoil extends from a root at a blade platform, which defines the radially inner flowpath for the combustion gas, to a radially outer tip cap, and includes opposite pressure and suction sides extending axially from leading to trailing edges of the airfoil. The cooling circuit extends inside the airfoil between the pressure and suction sides and is bounded at its top by the airfoil tip cap. A squealer tip blade has a squealer tip wall extending radially outwardly from the top of the tip cap and around the perimeter of the airfoil on the tip cap to define a radially outwardly open tip cavity.
The squealer tip is a short radial extension of the airfoil wall and is spaced radially closely adjacent to an outer turbine shroud to provide a relatively small clearance gap therebetween for gas flowpath sealing purposes. Differential thermal expansion between the blade and the shroud, centrifugal loading, and radial accelerations cause the squealer tips to rub against the turbine shroud and abrade. Since the squealer tips extend radially above the tip cap, the tip cap itself and the remainder of the airfoil is protected from damage, which maintains integrity of the turbine blade and the cooling circuit therein.
However, since the squealer tips are solid metal projections of the airfoil, they are directly heated by the combustion gas which flows thereover. They are cooled by heat conduction with the heat then being removed by convection into the tip cap and cooling air injected into the cavity by passages through the tip. The cooling air from within the airfoil cooling circuit is used to convect heat away from tip and to inject into cavity. The squealer tip typically operates at temperatures above that of the remainder of the airfoil and can be a life limiting element of the airfoil in a hot turbine environment.
Thermal barrier coatings (TBC) are well known and proven as thermal insulators used at various locations in gas turbine engines. However, TBC is effective only at locations in the engine where heat flux is high due to differential temperature between hot and cold sides of a component. Since a typical squealer tip is directly bathed on both its inboard and outboard sides in the hot turbine flowpath gas, it has a relatively low heat flux laterally therethrough which decreases the effectiveness of TBC applied on the outboard side thereof.
Since the pressure side of an airfoil typically experiences the highest heat load from the combustion gas, a row of conventional film cooling holes is typically provided in the pressure side of the airfoil outer wall immediately below the tip cap for providing a cooling film which flows upwardly over the pressure side of the squealer tip. Although this enhances cooling of the pressure side squealer tip, it also effects a relatively large radial temperature gradient from the top of the squealer tip down to the tip cap near the film cooling holes. A large temperature gradient in this direction creates thermal stress which over repeated cycles of operation of the engine may lead to metal cracking that limits the effective life of the blade.
In order to reduce this undesirable radial thermal gradient in the squealer tips, the blade tips have been masked during the TBC coating process to eliminate TBC along the outboard side of the squealer tip, while maintaining TBC over the remainder of the outer surface of the outer wall of the airfoil. The entire squealer tip, in such a blade, is operated without TBC protection to reduce the undesirable radial temperature gradient. However, the masking process in the manufacture of the turbine blades significantly increases the cost of manufacture which is undesirable.
U.S. Pat. No. 5,733,102, entitled "Slot Cooled Blade Tip", discloses a slot extending radially inwardly to the tip cap and along the pressure squealer tip between leading and trailing edges of the airfoil. A plurality of spaced apart supply holes extend radially through the tip cap from the slot to the cooling circuit for channeling the coolant into the slot for cooling the squealer tip. A thermal barrier coating is disposed on an outboard side of the squealer tip for providing insulation against the hot gas that flows therealong. The construction of the turbine blade squealer tip in U.S. Pat. No. 5,733,102, is to eliminate the masking process, while still providing effective cooling of blade squealer tips, when used in conjunction with TBC.
TBC has not been used inside the tip cap cavities of rotating airfoils because of concerns that the thermal gradient from the top of the squealer tip to the tip cap area will be increased (cooler tip cap) which in turn would cause an increase in the stresses that generate commonly occurring squealer tip cracks. Squealer tip wall cracks occur due to operational environment and it is desirable to prevent them from propagating into the tip cap and also to lower tip cap operating temperature to improve material properties. The squealer tip cracking eventually begins to propagate across the tip cap or plenum. Several tip cap cracks propagate and join together in the tip cap resulting in the liberation of a portion of the tip cap. The missing tip cap portion "short circuits" the airfoil cooling circuit resulting in premature distress to the area of the airfoil receiving little to no cooling air. The squealer tip and tip cap cracking would most likely cause more complicated weld repairs to be performed at the blade service shops. These more complicated weld repairs result in increased losses at the engine overhaul level and more expensive blade repairs which both adversely impact the maintenance costs per flight hour of the engine. It is desirable to prevent tip crack propagation and avoid these costly weld repairs.